1. Field of The Invention
The invention concerns controlling the attitude of a satellite about its three axes, usually a geostationary satellite stabilized about the three axes, with optional compensation of disturbing torques acting on the satellite about the roll/yaw axes during orbit control maneuvers.
It is also concerned with the general configuration of satellites whose attitude is stabilized about their three geometrical axes in geostationary orbit, whether for civil or military, commercial or scientific purposes, or combinations thereof.
2. Description of the Prior Art
In the present context a satellite is any artificial object in the solar system in orbit around the Earth or any other planet or object of the solar system, or, in solar orbit, possibly a transfer orbit between two planets.
Satellites in orbit are known to be subjected to disturbing torques which make it necessary to control their attitude. The most important causes of disturbing torques are the lack of symmetry about the center of gravity (in space the expression "center of mass" is more appropriate) of the effects of solar radiation pressure (solar pressure for short) due to the angle of incidence (not 90.degree.) of the pitch axis of the satellite to the Sun, different reflectivity characteristics of parts of the satellite and geometrical asymmetries of the satellite; the action of the local (for example terrestrial) magnetic field; the aerodynamic effect of the environment (in low orbits); and the distance from the center of gravity of the satellite to the resultant thrust vector axis of the thrusters used to modify the Satellite orbit.
It is possible to distinguish between disturbances related to the environment: solar pressure, interaction of the satellite magnetic dipole with the surrounding magnetic field, gravitational gradient, etc.; these disturbances are weak (order of magnitude=10.sup.-5 N.m) but act on the satellite at all times, and disturbances related to misalignment of the thrust vector of the orbit control chemical thrusters relative to the center of gravity of the satellite; these disturbances are stronger (order of magnitude=10 N.m.s per day for a geostationary satellite) but limited in time.
It is essential to provide means for controlling the attitude of a satellite in its orbit. Various active means have already been proposed for this purpose, using a plurality of reaction wheels or thrusters of the mass ejection type, but the principle of ejecting mass requires that the satellite carry a reserve of mass, which increases the weight of the satellite at launch. Further, gas jet thrusters cause intense disturbances which excite the flexible and nutation modes of the satellite, degrading pointing accuracy. Additionally, low-thrust type thrusters such as ion thrusters or electrical arc ionization thrusters consume considerable electrical power and require warm-up phases which generally lead those skilled in the art to avoid using them for attitude control, and reaction wheels are not sufficient in themselves because the wheels must be desaturated to bring their speed to a value near the nominal value from time to time and this requires the application of external torque to the satellite.
To control the attitude of a satellite for an optimal mass budget use is made of disturbing forces due to solar pressure, by appropriately orienting surfaces attached to the satellite, or the local, for example terrestrial, magnetic field, by creating magnetic dipoles on board the satellite by means of pairs of currents.
Various prior art references have already put forward the use of solar radiation pressure for satellite attitude control and orbit control (stationkeeping) using mobile surfaces which can be oriented by means of dedicated actuators or using orientation thrusters already on board.
French Patent 2,513,589 describes a method and a device for aligning with a required direction the roll axis of a satellite which is spin-stabilized and fitted with a plurality of fixed solar panels; mobile surfaces are mounted at the ends of the panels.
French Patent 2,550,757 proposes to control the position of satellites by acting on the solar panels by deforming them to impose a variable backwards curvature on each of them.
French Patent 2,529,166 concerns a satellite stationkeeping method using solar sails and a space vehicle implementing this method. Solar sails disposed to the North and South are mounted on the satellite at the end of pylons parallel to the North-South axis. The pylons can rotate on themselves and the sails can be inclined about axes transverse to the pylons.
German Patent 2,537,577 entitled "Satellite Attitude Control", teaches the provision at the end of the solar panels of surfaces that can be oriented about the axis of the solar panels and transversely thereto.
U.S. Pat. No. 3,304,028 entitled "Attitude Control for Spacecraft", is similar to French Patent 2,513,589, previously mentioned, as is U.S. Pat. No. 3,339,863.
French Patent 2,530,046 entitled "Geosynchronous Satellite Attitude Control Method and Device", teaches the addition of fixed surfaces to the sides of the solar panels.
French Patent 2,531,547 entitled "Geostationary Satellite Attitude Control System", teaches variation of the relative orientations of the solar panels about their axes as does U.S. Pat. No. 4,325,124 entitled "System for Controlling the Direction of the Momentum Vector of a Geosynchronous Satellite".
European Patent 0,295,978 proposes a device and a method for pointing a space probe towards a heavenly body. North and South solar sails are added to the satellite which have asymmetrical surface areas, orientations about a North-South axis or inclinations transverse to this axis.
French Patent 2,552,614 proposes a Satellite configuration with improved solar means comprising solar panels oriented transversely to the North-South axis and adapted to be oriented about axes transverse-to the North-South axis.
Finally, U.S. Pat. No. 4,262,867 provides for solar panels adapted to be partially retracted accordion fashion to each side of which solar sails are hinged about axes transverse to the longitudinal axis of the panels.
These prior art references concern attitude control devices which use solar pressure as their means of actuation. However, all the solutions taught therein have one or other of the following drawbacks. Either they require extra surfaces to be added, with the disadvantages that the additional surfaces increase the mass of the satellite; the addition of mechanisms dedicated to deploying the surfaces in orbit increases the mass and the risk of failure; and the additional overall dimensions due to the surfaces represent a satellite volume penalty at launch; or they provide satellite attitude control about one or two axes only, requiring further means for control about the third axis.
Various documents have proposed displacing the center of gravity of the satellite to reduce the disturbing torques related to misalignment between the center of gravity and the thruster (or solar pressure) thrust vector. They include U.S. Pat. No. 4,684,084 entitled "Spacecraft Structure with Symmetrical Mass Center and Asymmetrical Deployable Appendages"; U.S. Pat. No. 4,345,728 entitled "Method for Controlling the Attitude of a Spinning Spacecraft in the Presence of Solar Pressure"; and U.S. Pat. No. 3,516,623 entitled "Stationkeeping System".
U.S. Pat. Nos. 3,516,623 and 4,345,728 propose reducing the disturbing torques acting on a spin-stabilized satellite related to misalignment of the center of gravity and the thrust by moving the center of gravity using mobile weights, these weights and their actuators having no other function.
U.S. Pat. No. 4,684,084 describes a satellite configuration in which the disturbing torques due to misalignment between the center of gravity and the thrust vector of the orbit control thrusters are reduced. The center of gravity is moved towards the thrust axis by appropriate positioning of the solar generator panels after they are deployed. This positioning is fixed and not variable in flight. This configuration is such that the center of gravity is substantially fixed despite the deployment of highly asymmetric appendages, but there is no possibility of modifying the position of the solar generator panels in flight. This has the drawback of increasing the disturbing torques due to solar radiation and of providing nothing to compensate this.
As for the propulsion employed during the operational phase of current three-axis stabilized satellites, in particular in the United States, Japan and Europe, this is purely chemical (using hydrazine or a mixture of propellants, for example) or chemical with electrical assistance (example: power augmented catalytic thruster (PACT) heated or electrical arc (Arcjet) hydrazine or ion or plasma thrusters for orbit correction.
However, in the final analysis attitude control is achieved by chemical propulsion with intermediate Biorage of angular momentum in one or more inertia wheels about two or three axes.
Relevant publications include: "The Attitude Determination and Control Subsystem of the Intelsat V Spacecraft"--Proceedings of the AOCS Conference, Noordwijk, October 1977; "Precision Attitude Control with a Single Body-Fixed Momentum Wheel"--AIAA Mechanics and Control Flight Conference--Anaheim, Calif., August 1974; U.S. Pat. No. 4,949,922 entitled "Satellite Control System"; and "Satellite Attitude and Orbit Control System: Developments to the 80-90's"--L'Aeronautique et l'Astronautique--no. 69, 1878-2--p 33-56.
Similarly, the use of electrical propulsion for orbit control and even attitude control is under widespread consideration at the present time, as indicated by the following publications, in particular, "Electric Propulsion Projects and Researches in Japan", AIAA 20th International Electric Propulsion Conference, Garmisch, Partenkirchen (Germany), October 1988; "Design and Integration of an Electric Propulsion System on the Eurostar Spacecraft", same conference as above; "Readiness Appraisal: Ion Propulsion for Communication Satellites", AIAA 12th International Communication Satellite Systems Conference, Crystal City, March 1988; and "Chemical and Electric Propulsion Tradeoffs for Communication Satellites", Comsat Technical Review Volume 2 Number 1, Spring 1972, pp 123-145.
With reference to ion thruster propulsion as such, reference may be made to French Patent 2,510,304 entitled "Field Emission Ion Source Suitable for Electric Propulsion of Space Craft"; U.S. Pat. No. 3,279,176 entitled "Ion Rocket Engine"; and U.S. Pat. No. 4,829,784 entitled "Method and System for Storing Inert Gas for Electrical Impulse Space Drives".